Aircraft manufacturers continuously attempt to improve aircraft performance by reducing both weight and manufacturing costs while maintaining or improving structural strength. One well-known method for increasing aircraft performance is to reduce airframe weight through the use of state-of-the-art materials, such as composites, having relatively high strength-to-weight and stiffness-to-weight ratios. Composite materials are generally described as being materials that include reinforcing fibers, such as graphite fibers, embedded in a polymeric matrix, such as an epoxy resin. Such materials will hereinafter be referenced as “fiber-reinforced composite” materials. Fiber-reinforced composite materials are usually supplied as fibrous sheets pre-impregnated with a curable or partially cured resin. The so-called “prepreg sheets” may then be laid up in laminated plies and cured to form rigid panel structures.
Integrated composite structures which comprise elongate stringers or other structural reinforcement members integrated to a skin panel are also being employed in the aircraft industry. Typically, the uncured prepreg sheets forming the stringers are laid up in plies (usually cross-lapped) onto similarly uncured prepreg plies forming a panel preform positioned on suitable support tooling. Once the stringer layers are laid up, suitable mandrel structures may be positioned so as to assist in maintaining the structural form during the subsequent cure process.
The preformed panel and stringer are then typically covered by a removable flexible caul sheet to form a curing assembly. A vacuum bag can then be positioned over the curing assembly with suitable seals placed between the bag and the mold tooling. Curing of the panel and stringer can then be performed at an elevated temperature and typically elevated pressure in an autoclave or oven. After curing all prepreg plies, the vacuum bag and the flexible caul can be removed thereby providing a cured integrated composite panel and stringer component.
One problem associated with the use of a flexible caul during the curing of integrated prepreg layers is that there may sometimes occur non-conforming cure at the edges of the stringer or stiffener adjacent the panel due to improper edge ply compaction. It is towards addressing such problem that the embodiments of the present invention as described herein are directed.